Long endurance vertical takeoff and landing aircraft

ABSTRACT

An aircraft for use in fixed wing flight mode and rotor flight mode is provided. The aircraft can include a fuselage, wings, and a plurality of engines. The fuselage can comprise a wing attachment region further comprising a rotating support. A rotating section can comprise a rotating support and the wings, with a plurality of engines attached to the rotating section. In a rotor flight mode, the rotating section can rotate around a longitudinal axis of the fuselage providing lift for the aircraft similar to the rotor of a helicopter. In a fixed wing flight mode, the rotating section does not rotate around a longitudinal axis of the fuselage, providing lift for the aircraft similar to a conventional airplane. The same engines that provide torque to power the rotor in rotor flight mode also power the aircraft in fixed wing flight mode.

CROSS REFERENCE TO RELATED APPLICATIONS

This application is a continuation application of U.S. patentapplication Ser. No. 13/429,156, filed Mar. 23, 2012, which claims thebenefit of U.S. Pat. App. No. 61/465,760, filed Mar. 24, 2011. Thisapplication hereby incorporates by reference the above-identifiedapplications in their entirety.

FIELD OF THE INVENTION

The invention relates generally to aircraft designs, and, moreparticularly, to unmanned aircraft designs that combine the features offixed wing aircraft and vertical takeoff and landing (VTOL) aircraft.

BACKGROUND OF THE INVENTION

Tactical Unmanned Aerial Vehicles (UAVs) have revolutionized the waywars are fought and intelligence is gathered because of their low-cost,safety, and long endurance. However, one major limitation of currentUAVs is the need for prepared launch and recovery sites with largefootprints or a traditional airport. A portable system that does notrequire bulky launch equipment or runways would greatly increase theutility of manned and unmanned aircraft, much as helicopters do for lowendurance operations.

Launch and Recovery Equipment (LRE) for UAVs can be comprised of both alaunch vehicle as well as a recovery system, such as a large net. Thisequipment has many operational burdens. First, the LRE site must offer aclear launch and approach path. This path must be obstacle-free over amile or more beneath a 5 degree flight path, which makes urban basingvery difficult. Additionally, some UAVs need two dedicated techniciansper shift to handle LRE operations. LRE hardware typically weighs1,500-3,000 pounds for the launcher, and a similar amount for recoveryequipment. These units must be transported to the launch site, whichoften requires two or more Humvees for transport. Similarly, traditionallong range or high endurance manned aircraft need a runway with theattendant clearings below the takeoff and landing flight paths.

Previous aircraft designs attempt to combine the vertical takeoff andlanding (VTOL) and hover capabilities of a helicopter with the increasedspeed and range capabilities of fixed wing aircraft. These hybriddesigns reduce the footprint necessary for launch and recovery but aremore complex than either helicopters or conventional take-off andlanding aircraft as they generally incorporate multiple propulsionsystems, each used for a different flight mode. These designs caninclude “tail sitter” configurations, so named because the aircrafttakes off and lands from a tail-down orientation. Other designs caninclude “nose sitter” configurations, so named because the aircrafttakes off and lands from a nose-down orientation.

One example of a nose-sitter design includes a VTOL hybrid, whichincludes a conventional propeller for fixed wing flight and a foldingrotor near the tail of the aircraft. These designs may have high hoverefficiency; however, they also require complex mechanical systems andweigh more than other designs due to the requirement of two separatepropulsion systems, one for each flight mode.

Other VTOL designs can include “tail sitter” configurations, so namedbecause the aircraft takes off and lands from a tail-down orientation.Conversion from vertical to horizontal flight for these hybrid designstypically requires a configuration change and dedicated engines for eachconfiguration. Prior solutions that combine VTOL and cruise performancecompromise performance in both flight modes.

A VTOL airplane or UAV that uses the same propulsion for both flightmodes would have many structural benefits, including reduced complexityand weight of the launch equipment and ease of operation in more remotelocations, as well as numerous mission benefits that are enjoyed todayby helicopters. These include hover-and-stare in urban-canyons andsit-and-stare for extended silent surveillance. Further, sit-and-waitoperation allows the airplane or UAV to be pre-deployed to a forwardarea awaiting mission orders for remote launch of the aircraft. Uponreceiving the mission order, the vehicle can launch without leaving anyexpensive launch equipment at the launch site.

Some existing VTOL designs suffer from poor endurance and speed. Forwardflight efficiency may be improved by partial conversion to an aircraftlike the V-22 but endurance issues remain. Many VTOL aircraft alsorequire a high power-to-weight ratio. These aircraft may be used forhigh-speed flight if the aircraft is fitted with a special transmissionand propulsion system. However, achieving high endurance requiresefficiency at very low power. Thus the challenge exists to create avirtual gearbox that equalizes power and RPM for VTOL and fixed wingflight achieving highly efficient cruise with the benefits of a verticaltakeoff and landing configuration.

SUMMARY OF THE INVENTION

In at least one embodiment, the present application relates to anaircraft that overcomes the shortcomings of the prior art noted above.

In some embodiments, an aircraft and, in particular, an aircraft capableof fixed wing and rotor flight modes is disclosed. A fixed wing flightmode is defined as flight in which a fuselage is substantially parallelto the ground and a rotor flight mode is defined as flight in which afuselage is substantially perpendicular to the ground. The aircraftcomprises a plurality of wings, each wing having a spar such that eachwing is rotatable about the spar. The aircraft also comprises at leastone actuator coupled to the spar of each of the plurality of wings. Theaircraft further comprises a plurality of engines secured to said wings.The aircraft also comprises a fuselage further comprising a rotatingsupport rotatable about a longitudinal axis of the fuselage wherein saidplurality of wings are secured to the rotating support such that thewings rotate about a longitudinal axis of the fuselage.

In some embodiments, an aircraft and, in particular, an aircraft capableof fixed wing and rotor flight modes is disclosed. The aircraftcomprises a first wing having a spar with a first engine secured at aposition between 20% and 75% of semi-span of said first wing and asecond wing having a spar with a second engine secured at a positionbetween 20% and 75% of semi-span of said second wing. At least oneactuator is coupled to the spar of each of the first and second wings.The aircraft further comprises a fuselage further comprising a rotatingsupport rotatable about a longitudinal axis of the fuselage wherein saidfirst and second wings rotate with respect to said rotating support suchthat the actuator coupled to the spar of said first wing rotates saidfirst wing in a first direction around the span-wise axis of the firstwing and the actuator coupled to the spar of said second wing rotatessaid second wing in a second direction around the span-wise axis of thesecond wing such that said first and second engines are directed indifferent directions causing said first and second wings to rotate abouta longitudinal axis of the fuselage due to the thrust provided by eachof said engines.

In some embodiments, a method for changing the configuration of anaircraft between flight modes is disclosed. The method is achievedthrough changing the rate of motion of a rotating portion of a fuselagefrom a first flight mode to a second flight mode and rotating each of aplurality of wings in opposite directions.

In some embodiments, an aircraft and, in particular, an aircraft capableof fixed wing and rotor flight modes is disclosed. The aircraftcomprises a fuselage main body defining a longitudinal axis, a rotatingsection connected said fuselage main body, and a plurality of engines.The rotating section further comprises a plurality of wings, each winghaving a spar such that each wing is rotatable about said spar and atleast one actuator coupled to the spar of each of the plurality ofwings. The plurality of engines is attached to the rotating section. Therotating section is rotatable around the longitudinal axis.

BRIEF DESCRIPTION OF THE DRAWINGS

These and other features, aspects, and advantages of the presentinvention will now be described in connection with a preferredembodiment of the present invention, in reference to the accompanyingdrawings. The illustrated embodiments, however, are merely examples andare not intended to limit the invention.

FIG. 1 is a perspective view of an aircraft according to a preferredembodiment of the invention that converts between rotor flight mode andfixed wing flight mode, shown in fixed wing flight mode.

FIG. 2 is a front perspective view of the aircraft of FIG. 1, shown infixed wing flight mode.

FIG. 3a is a top view of the aircraft of FIG. 1, shown in fixed wingflight mode.

FIG. 3b is an enlarged view of volume 300 depicted in FIG. 3 a.

FIG. 4 is a rear perspective view of the aircraft of FIG. 1, shown infixed wing flight mode.

FIG. 5 is a perspective view of the aircraft of FIG. 1, shown in rotorflight mode prior to launch or after landing on the ground.

FIG. 6 is a side view of the aircraft of FIG. 1, shown in rotor flightmode prior to launch or after landing on the ground.

FIG. 7 is a side view of the aircraft of FIG. 1, shown in rotor flightmode with the wings having rotated about a longitudinal axis of thefuselage.

FIG. 8 is a top view of the aircraft of FIG. 1, shown in rotor flightmode.

FIG. 9 is a graphical and pictorial representation of a preferred methodof converting an aircraft between a rotor flight mode and configurationto a fixed wing flight mode and configuration.

FIG. 10 is a graphical and pictorial representation of a preferredmethod of converting an aircraft between a fixed wing flight mode andconfiguration and a rotor flight mode and configuration.

FIG. 11 is a top view of an aircraft with the engines and propellers ina “pusher” style configuration in which the propellers push the aircraftthrough the air rather than pull the aircraft.

FIG. 12 is a view of another embodiment of the aircraft having jetengines.

DETAILED DESCRIPTION OF EXEMPLARY EMBODIMENTS

The following detailed description is directed to certain specificembodiments of the invention. However, the invention may be embodied ina multitude of different ways as defined and covered by the claims.

Embodiments of the invention can provide the features of both verticallyflying, or hovering, aircraft and fixed wing aircraft. An aircraft thatis changing altitude while traveling in a substantially verticaldirection with the fuselage substantially perpendicular to the ground isdefined to be in rotor flight mode. The same aircraft flyinghorizontally with the fuselage substantially parallel to the ground isdefined as flying in a fixed wing flight mode. Some embodiments of theinvention desirably result in increased efficiency and long range flightendurance.

A preferred embodiment of the invention combines the wing function of afixed wing aircraft and the rotor function similar to that provided by atraditional helicopter rotor. In a preferred embodiment, when theaircraft is in rotor flight mode, the wings rotate around the fuselage,the rotation of the wings propelled by the same engines and propellersthat are used for cruise as an airplane. The rotation of the wings actssimilarly to the rotor of a traditional helicopter, providing verticalthrust to vertically propel the aircraft and maintain a hoveringaltitude. In the same preferred embodiment, when the aircraft is infixed wing mode, the wings do not rotate around the fuselage and thesame engines that provided torque to rotate the wings when the aircraftwas in hover mode provide the thrust required to power the aircraft infixed wing flight. This arrangement provides the features of a rotorflight, aircraft and a fixed wing aircraft while reducing performancelosses due to the weight requirements of complex mechanical machineryneeded for configuration changes. The preferred embodiment alsoeliminates the need for the aircraft to carry multiple propulsionsystems for flight in more than one flight mode. Embodiments of theinvention can include features such as but not limited to improvedpayload capacity, vertical take off and landing (VTOL) capability,efficient hover, high speed, and long range endurance in a singleflight. While the preferred embodiment discussed below includes engineswith propellers, other embodiments may include jet engines, as picturedin FIG. 12.

FIG. 1 depicts a preferred embodiment of the invention. In FIG. 1 theaircraft 100 is shown in fixed wing flight mode, similar to that of aconventional airplane, such as a Piper Seneca or Beech King-Air. Theaircraft 100 comprises a fuselage main body 102 having nose 104, payloadcompartment 106, wing attachment region 108, and tail section 110. Thewing attachment region 108 further comprises a rotating support orportion 502. The wings 112, 114 may be attached to rotating support 502.The fuselage 102 may be shaped as a cylinder with tapered ends or may beof other shapes known by those skilled in the art.

The payload compartment 106 may be located within the fuselage 102between the nose 104 and wing attachment region 108. For example, in theillustrated embodiment the payload compartment 106 may have a volume of250 cubic inches. Alternative embodiments may have a payload compartment106 volume between 50 cubic inches and 10,000 cubic inches, desirablybetween 100 cubic inches and 5,000 cubic inches, more desirably between150 cubic inches and 1,000 cubic inches, and even more desirably between200 cubic inches and 800 cubic inches. However, the payload compartment106 may be smaller or larger depending on the overall size of theaircraft and mission requirements.

FIG. 3 a shows that the interior of the fuselage 102 may also comprise avolume 300, shown enlarged in FIG. 3 b, which contains a flap and lagbearing, a centrifugal bearing, servos to control wing surfaces,actuators to control wing rotation, batteries, and power regulationequipment, as well as fuel tanks or other mission specific equipment.The interior equipment may be placed such that it does not interferewith the rotation of the wing attachment region 108 of the fuselage,discussed below.

As seen in FIGS. 1-8, the wings 112, 114 are attached to the wingattachment region 108 of the fuselage 102 with one wing on each side ofthe fuselage. As pictured in FIG. 4, the wing attachment region 108comprises a rotating support 502 to which the wings 112, 114 preferablyattach. The wings 112, 114 and rotating support 502 rotate around alongitudinal or lengthwise axis of the fuselage 102 when the aircraft100 is in rotor flight mode. The rotating support 502 is preferablylocked to the fuselage 102 to prevent rotation about a longitudinal axisof the fuselage when the aircraft is in fixed wing flight mode. The wingattachment region 108 may desirably comprise a length of the fuselageequal to the chord length of the wings at the point of attachment withthe wing attachment region 108, or it may be of lesser or greaterlength.

As pictured in FIG. 1, the wings 112, 114 may also comprise one or morecontrol surfaces 120, 122 to control the attitude of the aircraft whilein both fixed wing and in rotor flight modes. These control surfaces maybe controlled by servos located within the fuselage 102 of the aircraft100.

In a preferred embodiment the wings may comprise a symmetric airfoil.The wings 112, 114 each have a leading edge 124, 126, and a trailingedge 128, 130. The wings 112, 114 preferably have a greater chordlength, or leading edge 124, 126 to trailing edge 128, 130, closer tothe fuselage 102, as shown in FIG. 1, or the wings 112, 114 may havesubstantially the same chord length along the span of the wing from wingtip 116 to wing tip 118.

In FIG. 1 one engine 132, 134 is secured to each wing 112, 114. In otherembodiments, one or more engines may be secured to each wing 112, 114. Arotating section includes the wings 112, 114 and the rotating support502. The engines 132, 134 with propellers 136, 138 are connected to therotating section. In a preferred embodiment, the engines 132, 134 of theaircraft 100 are aligned substantially parallel with a longitudinal axisof the fuselage with the propellers 136, 138 configured to pull theaircraft 100 through the air when the aircraft 100 is in fixed wingflight, as depicted in FIG. 2. In this flight mode, the rotatingsection, comprised of wings 112, 114 and rotating support 502 withengines 132, 134 and propellers 136, 138 attached to the rotatingsection, does not rotate about a longitudinal axis of the fuselage 102.In other embodiments, the engines 132, 134 and propellers 136, 138 maybe configured in a push-type configuration in which the propellers 136,138 are oriented toward the tailing edge 128, 130 of the wings 112, 114to push the aircraft 100 rather than to pull the aircraft 100 when theaircraft 100 is flying in a fixed wing flight mode. A “pusher” styleconfiguration where the engines and propellers are oriented to push theaircraft 100 through the air is shown in FIG. 11.

The fuselage 102 may also comprise a tail section 110, as depicted inFIG. 1. The tail section 110 may be comprised of three or morestabilizing surfaces 140, 142, 144, 146 extending from the fuselage 102in approximately equal intervals about the circumference of the tailsection 110. In a preferred embodiment shown most clearly in FIGS. 3 and4, four tail stabilizing surfaces are shown, though other embodimentsmay have a different number of tail stabilizing surfaces. A plurality oftail stabilizing surfaces 140, 142, 144, 146 may incorporate a controlsurface 148, 150, 152, 154 such as a rudder or ruddervator surface tocontrol the attitude of the aircraft while in flight. The controlsurfaces 148, 150, 152, 154 may be controlled by servos located withinthe tail section 110 of the fuselage 102. As shown in FIG. 1, tailstabilizing surfaces 140 and 144 comprise control surfaces 148, 152;however, the other tail surfaces 142 and 146 may also comprise controlsurfaces. In a preferred embodiment, each tail stabilizing surface 140,142, 144, 146 has a control surface 148, 150, 152, 154. The controlsurfaces on tail stabilizing surfaces 142, 146 are not shown in FIG. 1.Each control surface 148, 150, 152, 154 can act independently ortogether with the other control surfaces as a rudder or ruddervator tocontrol the direction of the aircraft's flight in the yaw direction. Theyaw direction is defined as rotation about an axis perpendicular to thehorizontal plane P of the aircraft defined by the wings when in a fixedwing flight mode and a longitudinal axis of the fuselage. In otherembodiments, more or fewer tail surfaces may comprise control surfacesfor controlling the attitude of the aircraft 100.

Each tail stabilizing surface 140, 142, 144, 146 may also have one ormore landing pads 156, 158, 160, 162 to contact the ground and supportthe aircraft 100 while it is in a “tail-sitter” configuration at thelaunch site in preparation for launch or when the aircraft lands. FIGS.1 and 2 depict a preferred embodiment in which one landing pad 156, 158,160, 162 is located at the tip of each tail stabilizing surface 140,142, 144, 146 for contact with the ground to support the aircraft 100prior to and during launch and after landing.

FIG. 2 is a perspective view of the aircraft 100 from a nearly head-onposition when the aircraft is in a fixed wing flight mode for horizontalflight as an airplane. In this preferred embodiment, the wings 112, 114and rotating support 502 of the aircraft 100 may be locked with respectto the fuselage 102, meaning that they may not be permitted to rotateabout a longitudinal axis of the fuselage 102. In FIG. 2, each wing 112,114 has been rotated such that the propeller 136, 138 of each engine132, 134 is located substantially in the desired direction of travel.

The embodiment depicted in FIG. 2 also illustrates that the engines 132,134 are secured to the wings 112, 114 and are not located within thefuselage 102. Locating the engines 132, 134 on the wings 112, 114 mayreduce or eliminate the need for extension shafts in a preferredembodiment. Extension shafts typically connect an engine mounted withinor directly on the fuselage via a gearbox or other linkage to thepropellers on the wing. Locating the engines within or directly on thefuselage typically also requires a central gearbox located within thefuselage. By eliminating the extension shafts and the central gearbox ina preferred embodiment, the weight of the aircraft 100 may be decreased,allowing for greater payload capacity, longer range, and endurance,among other benefits conceivable by those skilled in the art.

However, in other embodiments, engines 132, 134 may be secured at anypoint on the rotating section comprising the wings 112, 114 and rotatingsupport 502. In the illustrated embodiment, two engines are depicted.Additional embodiments may have more or fewer numbers of enginesdepending on mission requirements, other aircraft design considerations,or other considerations known to those skilled in the art.

FIG. 2 further illustrates that, in a preferred embodiment, the engines132, 134 are attached to the wings 112, 114 at a position an equaldistance to either side of the fuselage 102. Locating the engines 132,134 in this balanced orientation may provide benefits of balance andstability to the aircraft. Additionally, the engines 132, 134 arepreferably secured to the wings 112, 114 at an equalizing position alongthe semi-span of each wing, defined as the distance along the wing 112or 114 from the wing attachment region 108 to the wing tip 116 or 118.When the engines 132, 134 are located at the equalizing position in apreferred embodiment, the thrust of the engines 132, 134 and the flightspeed of aircraft 100 when the aircraft 100 is flying in a fixed wingflight mode desirably equal the torque and rpm, or rotations per minute,required by the aircraft 100 when the wings rotate around a longitudinalaxis of the fuselage 102 when the aircraft 100 is operating in a rotorflight mode. In a preferred embodiment, the torque demands of the wings112, 114 when acting as a rotor are matched to the in-flight demands ofthe aircraft 100 when flying in fixed wing mode, using the same engines132, 134 and propellers 136, 138. Locating the engines 132, 134 at thepoint where these demands are matched may also allow the wing tip 116,118 speed to approach sonic (when the wings 112, 114 are acting as arotor in rotor flight mode) while keeping the blades of the propellers136, 138 well under sonic. Locating the engines 132, 134 at the pointwhere these forces and requirements equalize preferably eliminates theneed for complex gearboxes and other heavy equipment that may decreasethe long range endurance capabilities of the aircraft. Additionaldiscussion of the determination of this point where these forces andrequirements equalize is included below.

In a preferred embodiment, the aircraft 100 may have multiple tailsurfaces 140, 142, 144, 146 to control the aircraft's attitude while inflight. In FIG. 2, the aircraft 100 is depicted in a preferredembodiment with four tail surfaces 140, 142, 144, 146. FIG. 2 depicts ahorizontal plane P defined by an axis of symmetry of wings 112, 114 anda longitudinal axis of the fuselage 102. In the pictured embodiment, twotail surfaces 140, 142 appear above the horizontal plane P of theaircraft and two of the tail surfaces 144, 146 appear below thehorizontal plane P. The spacing of the tail surfaces 140, 142, 144, 146around the tail section 110 of the fuselage 102 may provide stabilityfor the aircraft 100 when it is resting on the landing pads 158, 160,162, 164 located on the ends of each tail surface 140, 142, 144, 146prior to and during launch or after landing. The landing pads 158, 160,162, 164 prevent damage to the tail stabilizing surfaces 140, 142, 144,146 by providing a stable support upon which the aircraft 100 can restprior to and during launch and after landing. The aircraft 100 may becomprised of more or fewer tail stabilizing surfaces 140, 142, 144, 146which may or may not be equipped with landing pads 158, 160, 162, 164,depending on the configuration of the aircraft and launch and landingrequirements of the aircraft.

FIG. 3 a depicts an overhead view of the aircraft 100 in fixed wingflight mode. In this flight mode, the leading edges 124, 126 of thewings 112, 114 preferably face substantially in the same direction as alongitudinal axis of the fuselage. The engines 132, 134 and propellers136, 138 also face in the same direction as a longitudinal axis of thefuselage and provide the necessary power to propel the aircraft 100through the air in flight similar to that of a conventional airplane.The wings 112, 114 are preferably attached to the fuselage 102 viarotating support 502 located within the wing attachment region 108.Rotating support 502 is a section of the fuselage 102 that has theability to rotate around a longitudinal axis of the fuselage 102. Therotating support 502 may be locked, preventing rotation with respect tothe fuselage 102, but such locking is not required. As shown in FIG. 3a, at least one landing pad 158, 160 may be located at the tip of eachtail stabilizing surface 140, 142 for contact with the ground to supportthe aircraft prior to and during launch and after landing. The interiorof the wings 112, 114 may also comprise one or more fuel tanks forsupplying the engines, as pictured in FIG. 3 a.

With continued reference to FIG. 3A, the wings 112, 114, each include arear swept portion 113, 115 proximate to the fuselage 102 and a forwardswept portion distal to the rear swept portion 117, 119. In theexemplary embodiment, the engines 132, 134 are between the rear sweptportion and the forward swept portion of the wings 112, 114,respectively.

A rear perspective view of a preferred embodiment of the aircraft infixed wing flight mode is depicted in FIG. 4. As described above withrespect to FIGS. 1-3, preferably the leading edges 124, 126 of the wings112, 114 point substantially in the direction of travel similar to aconventional airplane such a Piper Seneca or Beech King-Air. The wings112, 114 are attached to rotating support 502, which is part of wingattachment region 108 of the fuselage 102. In the fixed wing flight modepictured in FIG. 4, the wings preferably do not rotate about alongitudinal axis of the fuselage 102. The engines 132, 134 with theirattached propellers 136, 138 point in substantially the same direction,as for a conventional airplane, and provide thrust to power the aircraft100 in a fixed wing flight mode. In this embodiment, the tail section110 of the fuselage 102 comprises four tail stabilizing surfaces 140,142, 144, 146. As has been discussed above, other embodiments of theaircraft may comprise fewer or greater numbers of tail stabilizingsurfaces. FIG. 4 clearly depicts two of the tail stabilizing surfaces,specifically 142 and 146, which further comprise control surfaces 150,154; however, the other tail stabilizing surfaces 140 and 144 may alsocomprise control surfaces. As discussed above with respect to FIG. 1,these control surfaces may act independently or together to control theattitude of the aircraft 100 while in flight. These control surfaces maypreferably be controlled by servos located within the tail section 110of the fuselage 102.

For flight in rotor flight mode, each wing 112, 114 may preferably berotated nearly 90 degrees in opposite directions about their length orspan-wise axis, as illustrated in a preferred embodiment in FIG. 5, suchthat the engines 132, 134 face in different directions. However, inother embodiments, each wing 112, 114 could rotate between 30 and 100degrees in opposite directions, between 60 and 90 degrees in oppositedirections, and between 75 and 85 degrees in opposite directions. Insome embodiments, each wing 112, 114 rotates at least 90 degrees, whilein other embodiments each wing 112, 114 rotates at least 60 degrees,while in still further embodiments, each wing 112, 114 rotates at least30 degrees. In some embodiments, the wings 112, 114 may be rotated inopposite directions such that the included angle between them is between30 degrees and 180 degrees. However, in other embodiments the includedangle between wings 112, 114, and by extension, the included anglebetween engines 132 and 134, could be between 60 and 180 degrees,between 90 and 180 degrees, between 120 and 180 degrees or between 150and 180 degrees. In other embodiments, the included angle betweenengines 132 and 134 and wings 112 and 114 could be at least 30 degrees,at least 60 degrees, at least 90 degrees, at least 120 degrees, or atleast 150 degrees. In other embodiments, wing 112 and wing 114 may noteach rotate the same amount. In these embodiments, wing 112 may rotateforward or backward between 30 and 100 degrees, between 60 and 90degrees, or between 75 and 85 degrees. Correspondingly, wing 114 mayrotate backward or forward between 30 and 100 degrees, between 60 and 90degrees, or between 75 and 85 degrees. Preferably, wings 112 and 114rotate at least 30 degrees in opposite directions but in otherembodiments wings 112, 114 may rotate at least 60 degrees or at least 90degrees. FIG. 5 depicts one embodiment of the aircraft 100 after thewings 112, 114 have been rotated in opposite directions. The rotation ofthe wings 112, 114 may preferably be achieved by servos or actuatorslocated within the fuselage 102.

Additionally, engines 132, 134 may also rotate relative to the wings112, 114 around a span or lengthwise axis of the wings 112, 114. Therotation of engines 132, 134 around a spanwise axis of the wings 112,114 may be in addition to the rotation of wings 112, 114 describedabove. The rotation of engines 132, 134 may be between 0 and 20 degrees,desirably between 0 and 10 degrees, or more desirably between 0 and 5degrees.

Preferably, the wings 112, 114 each have at least one spar. A spar runslengthwise along the internal or external span of the wing fromconnection with the fuselage to the wing tip to provide structuralrigidity. At least one spar of each wing 112, 114 attaches to therotating support 502 of wing attachment region 108 of the fuselage. FIG.5 depicts one spar 520 of wing 112 connected to rotating support 502 inwing attachment region 108. The wings may rotate about the spar or aspan-wise or wing tip to wing tip axis of the wing to position the wings112, 114 for hover or vertical flight. Desirably, spar 520 extends atleast to the point of attachment of engine 132 on wing 112 to providestructural rigidity to the wing 112. Wing 114 may be attached to wingattachment region 108 via a second spar, not shown in FIG. 5. Wing 112is preferably able to rotate as described above about the spar 520 toorient engine 132 and propeller 136 to a new direction required to powerrotation of wing 112 around a longitudinal axis of fuselage 102.Desirably, wing 114 also rotates about a second spar to achieve theorientation of engine 134 and propeller 138 as depicted in FIG. 5.

After rotation, the propellers 136, 138 of each engine 132, 134 face insubstantially opposite directions, as shown in a preferred embodiment inFIG. 5. Each wing 112, 114 is mounted to the rotating support 502located within wing attachment region 108 of the fuselage 102. Therotating support or portion 502 may be a section of the fuselage 102that comprises a solid disk or it may be a hollow member that may bebarrel shaped. The rotation of a rotating section, comprised of thewings 112, 114 and rotating support or portion 502 along with engines132, 134 attached to the rotating section, occurs as a result of thepower generated by the engines 132, 134. The power generated by engines132, 134 when the wings 112, 114 are in the preferred embodiment asshown in FIG. 5 results in the wings 112, 114 acting like the rotor of aconventional helicopter. The wings 112, 114 provide thrust in agenerally upward direction, causing the aircraft 100 to fly upwards in asubstantially vertical direction or to hover at a specified altitude.

As shown in FIGS. 5, 6, 7, and 8 the engines 132, 134 preferably areattached to the wings 112, 114 such that the rotating inflow speed ofair to the engines 132, 134 when the wings 112, 114 are acting as arotor is substantially similar to the cruise inflow speed of air to theengines 132, 134 when the aircraft 100 is flying in fixed wing mode.This preferably allows the propellers 136, 138 and the engines 132, 134of the aircraft 100 to be optimized for efficient cruise. The aircraft100 also relies on the same engines 132, 134 as those used for verticaltakeoff and landing and hovering flight when the aircraft 100 is infixed wing flight. In a preferred embodiment, there is notorque-to-ground force as is found with traditional helicopter designs,so no tail rotor is needed. Instead, the tail surfaces 140, 142, 146,148 are located in the rotorwash, defined as air driven downwards by therotation of the wings 112, 114 of the aircraft 100 around a longitudinalaxis of the fuselage 102 when the wings 112, 114 operate as a rotor forrotor flight. The location of the tail surfaces 140, 142, 144, 146 inthis rotorwash, plus small control surface 148, 150, 152, 154deflections, can cancel the small torque forces due to bearing drag thatact to rotate the fuselage 102 in the same direction of rotation as thewings 112, 114 when they act as a rotor. Desirably, after balancingrotor torque with throttle and rotor lift with pitch, (forces created bythe rotation of the wings 112, 114 around a longitudinal axis of thefuselage 102) the aircraft 100 ascends. Fuselage roll control isaffected by the interaction of the tail control surfaces 148, 150, 152,154 and the rotorwash generated by the rotation of the wings 112, 114about a longitudinal axis of the fuselage 102 when the wings 112, 114are acting as a rotor. Fuselage roll is defined as rotation about alongitudinal axis of the fuselage 102. The yaw directions is an axisperpendicular to a plane defined by a longitudinal axis of the fuselage102 and the span of the wings 112, 114 when the wings 112, 114 do notrotate relative to a longitudinal axis of the fuselage 102.

As shown in FIG. 5, takeoff and rotor flight is achieved when the wings112, 114 are preferably oriented substantially parallel to the groundwith the engines 132, 134 facing in opposite directions. FIG. 5 depictsone embodiment of the invention in which one engine 134, 136 is attachedto each wing 112, 114; however, a different number of engines may beattached to each wing. The application of power via the rotation of thepropellers 136, 138 attached to each engine 134, 136 causes the wings112, 114 to rotate around a longitudinal axis of the fuselage 102similar to a helicopter rotor in the direction indicated 1002 in FIG. 5.The pitch, or angle of attack, of each wing 112, 114 may be altered atthe same time (known in the art as collective pitch) or may be changeddepending on the position of each wing 112, 114 as it rotates (known inthe art as cyclic pitch). These pitch changes may be provided by controlsurfaces on the wings 112, 114 such as flaps, tabs with free-to-pitchwing bearings, or dedicated servos. As depicted in FIG. 5, the engines132, 134 are attached to the wings 112, 114 at a position where thetorque demands of the rotor created by the rotation of the wings 112,114 about a longitudinal axis of the fuselage 102 are matched to thein-flight demands of the aircraft 100 when the wings 112, 114 do notrotate relative to the fuselage in fixed wing flight mode. In apreferred embodiment such as that shown in FIGS. 1-10, the aircraft 100uses the same engines 132, 134 and propellers 136, 138 for flight infixed wing mode and rotor flight mode. This configuration may also allowthe rotor tips 116, 118 to approach sonic speed while keeping thepropellers 136, 138 well under sonic.

FIGS. 6 and 7 are side views of a preferred embodiment of the aircraft100 in a rotor flight mode configuration. As described above withrespect to FIG. 5, the wings 112, 114 are desirably rotated in oppositedirections along a span-wise axis of the wings 112, 114. In thepreferred embodiment pictured, the orientation of the wings 112, 114enables the thrust generated by the propellers 136, 138 to turn thewings 112, 114 about a longitudinal axis of the fuselage 102 in thedirection indicated 1002. For example, the same engines 132, 134 andpropellers 136, 138 that provide the thrust necessary to turn the wings112, 114 like a rotor when the aircraft 100 is in rotor flight mode alsoprovide between 50% and 100% of the thrust necessary to fly the aircraft100 in fixed wing flight mode, as depicted in FIGS. 1-4. In otherembodiments, engines 132, 134 desirably provide between 75% and 100% ofthe thrust necessary to fly the aircraft 100 in fixed wing flight mode,and more desirably provide between 90% and 100% of the thrust necessaryto fly the aircraft 100 in fixed wing flight mode as shown in FIGS. 1-4.In some embodiments, at least 50% of the thrust necessary to flyaircraft 100 in fixed wing flight mode is provided by the same engines132, 134 that power the aircraft in rotor flight mode, while in otherembodiments desirably at least 75% of the necessary thrust is providedby the same engines 132, 134, while in still other embodiments moredesirably at least 90% of the necessary thrust is provided by the sameengines 132, 134.

FIG. 8 is an overhead view of one embodiment of the aircraft 100 when itis in a rotor flight mode configuration. As discussed with respect toFIGS. 5-7 and as more clearly demonstrated here, each wing 112, 114desirably may be rotated in opposite directions about their length orspan-wise axis. In this configuration, the engines 132, 134 attached toeach wing 112, 114 may face in substantially opposite directions. As maybe more clearly seen in FIG. 8, each wing 112, 114 may comprise a spar802, 804 that runs lengthwise through the wing from the point ofattachment with the fuselage 102 to at least the point of attachment ofengine 132, 134 with wing 112, 114. Each spar 802, 804 providesstructural rigidity for each wing 112, 114, as may be appreciated bythose skilled in the art.

In a preferred embodiment, the spar 802, 804 of each wing is attached torotating support 502. As more clearly seen in FIG. 5, the wingattachment region 108 is a section of the fuselage 102 that comprises arotating support 502. A rotating section, comprised of rotating support502 and wings 112, 114 with engines 132, 134 attached to the rotatingsection, is allowed to rotate with respect to a longitudinal axis of thefuselage 102. The spars 802, 804 are preferably attached to the rotatingsupport 502 such that each wing 112, 114 is allowed to rotate about theaxis defined by the spar 802, 804 such that the leading edge of one wing124 and the leading edge of the other wing 126 face in substantiallyopposite directions, as shown in one embodiment in FIG. 8. The rotationof the wings 112, 114 about their spars 802, 804 will also result in theengines 132, 134 attached to each wing to face in substantially oppositedirections. Power generated by the engines 132, 134 will turn thepropellers 136, 138 which will produce thrust causing the rotation ofthe wings 112, 114 about a longitudinal axis of the fuselage 102 in thedirection indicated 1002 in FIG. 8. This rotation will cause the wings112, 114 to act as the rotor blades such as those of a helicopter,producing lift and desirably enabling the aircraft 100 to fly and hoverin a substantially vertical orientation.

As has been discussed with respect to FIGS. 1-8, aircraft 100 preferablytransitions from a VTOL or rotor flight mode to a fixed wing flight modefor flight similar to that of a conventional airplane. A preferredtransition to fixed wing flight is shown in FIG. 9. At initiationposition A, the aircraft is shown with the engines and propellersoriented in opposite directions as shown in FIGS. 5-8. The aircraft maybe on the ground G1 awaiting take off or may be hovering or flying inrotor flight mode above the ground G2. Between positions A and B, theaircraft preferably climbs to approximately 500 ft above ground level.At both positions A and B, the wings are rotating about the fuselage ofthe aircraft and acting as a rotor to provide thrust for rotor flight.At throttle down position B, the aircraft is preferably throttled downfrom a climb to hover while in rotor flight mode. Between throttle downposition B and pitch-over position C, the aircraft preferably begins apitch-over maneuver which transitions the attitude of the aircraft froma vertical orientation to a fixed wing orientation. During thepitch-over maneuver depicted between positions B and C, the wings of theaircraft are preferably rotated to a fixed wing flight position in whichthe engines face in substantially the same direction, this directionbeing the desired direction of travel for flight as a fixed wing orconventional airplane. For example, the transition of the wings from therotor flight mode to the fixed wing flight mode may occur within 1 to 15seconds, but in other embodiments the transition desirably may occurwithin 1 to 10 seconds or more desirably within 2 to 7 seconds. In someembodiments, the transition time is no more than 7 seconds or moredesirably no more than 5 seconds. The transition is accomplishedpreferably while simultaneously reducing engine throttle. The reductionin throttle desirably reduces rotor speed (the rotation of the wingsacting as a rotor) substantially to zero, which occurs betweenmid-transition position D and fixed wing flight mode position E. Atfixed wing flight mode position E, the aircraft has fully transitionedfrom a rotor flight mode to a fixed wing flight mode, meaning that thewings are no longer rotating around a longitudinal axis of the fuselagebut are substantially parallel to a centerline of the fuselage. Thewings may be locked with respect to the fuselage to prevent rotation butthis is not required. Additionally, the engines preferably facesubstantially in the direction of travel. At fixed wing flight modeposition E, engine throttle is preferably advanced, which acceleratesthe aircraft allowing for traditional fixed wing flight. Once sufficientairspeed is developed, the aircraft is flying “on-the-wing” similar tothat of a conventional airplane and may be controlled with conventionaltail surfaces.

FIG. 10 depicts a preferred method of transitioning from fixed wingflight to rotor flight. At fixed wing flight mode position E, theaircraft is oriented for flight in fixed wing mode, as described withrespect the same flight mode and position in FIG. 9. Between fixed wingflight mode position E and pitch-over position F, the aircraft desirablyis pitched over to accelerate the aircraft to greater than double thestall speed of the wings with the engines preferably at maximumthrottle. Between positions F and G, the aircraft then preferablyexecutes a pull-up maneuver to orient the aircraft for vertical flightwith the nose pointed vertically. This maneuver may generate forces thatact on the aircraft approximately between 2 and 4 times thegravitational force. At throttle reduction position G, the airspeed ofthe aircraft is reduced by throttling down the engines. Preferablybetween throttle reduction position G and rotor wing flight modeposition H, the aircraft's configuration is changed from that requiredfor fixed wing flight to that required for rotor flight. Thisconfiguration change preferably may occur within 1 to 15 seconds, but inother embodiments the configuration change desirably may occur within 1to 10 seconds or more desirably within 2 to 7 seconds, during which timethe wings rotate in opposite directions such that the engines andpropellers face in opposite directions as illustrated above in FIGS.5-8. In some embodiments, the time to change configurations is no morethan 7 seconds or more desirably no more than 5 seconds. At rotor wingflight mode position H, the wings begin to spin around a longitudinalaxis of the fuselage like the rotor of a helicopter due to the torquegenerated by the engines attached to the wings which now face inopposite directions. The rotor speed at rotor wing flight mode positionH is preferably increased beyond the speed required for hover flight.Finally, between rotor wing flight mode position H and fullytransitioned position J, the engines may be throttled down for stabledescent and landing. However, actual landing of the aircraft at thispoint may not be required if mission considerations and requirementsrequire the aircraft to maintain hover flight at a specific altitude orto complete other aerial maneuvers while in vertical flight mode.

As mentioned above with regard to FIG. 2, the torque demands of thewings 112, 114 when acting as a rotor are desirably matched to thein-flight demands of the aircraft 100 when flying in fixed wing mode,using the same engines 132, 134 and propellers 136, 138. The engines132, 134 are desirably positioned at a point on the wings 112, 114 wherethese requirements are substantially equalized. As discussed above,these requirements may have a difference between them of between 0% and50%, desirably between 0% and 25%, or more desirably between 0% and 10%.In some embodiments, the difference between these requirements isdesirably no more than 25% or more desirably no more than 10%. Thefollowing discussion describes a preferred method to calculate theposition on wings 112, 114 where the engines 132, 134 are attached tosubstantially equalize these requirements. The exact values used in thecalculation are for example purposes and are not intended to limit thecalculation or the invention in any way.

The table below provides a list of abbreviations used in the examplecalculations that follow:

VTOL Vertical Takeoff and Landing SHP Shaft horsepower (hp)PROP_Efficiency Propulsive Power/Input Power = Thrust * Vtrue/SHP at agiven flight condition GW Gross Weight (lbs) ROC Rate of Climb (feet perminute of fpm) Ceiling Maximum operating altitude of the airplane,typically defined as max power ROC = 100 fpm V and Vtrue True airspeed(feet per second or fps) V @ prop True airspeed at propeller station invertical flight mode (fps) VCruise True airspeed of aircraft in fixedwing flight mode (fps) ρ Air density (slugs/ft³) RPM Revolutions perminute (1/min.) L/D Fixed wing flight lift to drag ratio AR Wing aspectratio (wingspan²/wing area) CT${{Thrust}\mspace{14mu}{Coefficient}},{{{defined}\mspace{14mu}{as}\mspace{14mu}{CT}} = \frac{Thrust}{\rho*\left( \frac{R\; P\; M}{60} \right)^{2}*{Diameter}^{4}}}$Engine % Semispan Location of engine on semispan of wing, expressed as apercentage

It has been well established in the art that VTOL power required followsthis relation:

${VTOL\_ SHP}_{reqd} \cong \frac{\left( {RotorLift}_{reqd}^{1.5} \right)}{\left\lbrack {21*{RotorDiameter}*\sqrt{\frac{\pi}{4}}} \right\rbrack}$

Where VTOL_SHP_(reqd) is the Shaft horsepower required for vertical takeoff and landing.

Assuming that the aircraft requires 20% excess lift capability in therotor the equation for VTOL_SHPreqd becomes:

${VTOL\_ SHP}_{reqd} \cong \frac{\left( {1.2*G\; W} \right)^{1.5}}{\left\lbrack {21*{RotorDiameter}*\sqrt{\frac{\pi}{4}}} \right\rbrack}$

For an airplane the SHPreqd is set by the climb or takeoff requirementof the airplane. Since takeoff is not required when the aircraft is infixed wing flight mode, climb is the key consideration. Initial climbrate at takeoff altitude is a good surrogate for the ceiling capabilityof an airplane. The greater the ROC, or rate of climb, of an aircraft isat low altitude, the higher the ceiling, or the maximum altitude theaircraft may achieve. For many VTOL vehicles a typical ceiling is 15,000ft. This ceiling is approximately equivalent to a sea level ROC of 1,500fpm (or feet per minute) for a long range or high endurance airplane.Using the classical climb equation we can solve for the SHP requiredwhen the aircraft is climbing in fixed wing flight mode.

${CLIMB\_ SHP}_{reqd} = {\left( \frac{{VCruise}*G\; W}{{PROP\_ EFFICIENCY}*550} \right)*\left\lbrack {\frac{\left\lbrack \frac{R\; O\; C_{reqd}}{60} \right\rbrack}{V} + \left( \frac{L}{D} \right)^{- 1}} \right\rbrack}$

If the wings are used as the rotor, as in the preferred embodimentsdiscussed above with respect to FIGS. 1-8, the rotor diameter equals thewing span.

Further, if the flight engines are used to power the rotor, as discussedin the preferred embodiments pictured in FIGS. 5-8, the propellerefficiency must be included in the calculation to determine the engineSHP required for VTOL.

For VTOL, the equation becomes:

${ENGINE\_ SHP}_{reqd} \cong \frac{\left( {1.2*G\; W} \right)^{1.5}}{\left\lbrack {{PROP\_ Efficiency}*21*{RotorDiameter}*\sqrt{\frac{\pi}{4}}} \right\rbrack}$

For flight in fixed wing mode the equation becomes:

${ENGINE\_ SHP}_{reqd} = {\frac{{VCruise}*G\; W}{{PROP\_ Efficiency}*550}*\left\lbrack {\frac{\left\lbrack \frac{R\; O\; C_{reqd}}{60} \right\rbrack}{V} + \left( \frac{L}{D} \right)^{- 1}} \right\rbrack}$

Therefore;

$\frac{\left( {1.2*G\; W} \right)^{1.5}}{{PROP\_ Efficiency}*21*{RotorDiameter}*\sqrt{\frac{\pi}{4}}} = {\frac{{VCruise}*{GW}}{{PROP\_ Efficiency}*550}*\left\lbrack {\frac{\left\lbrack \frac{R\; O\; C_{reqd}}{60} \right\rbrack}{V} + \left( \frac{L}{D} \right)^{- 1}} \right\rbrack}$As an illustrative example only, for a very efficient 5000 lb airplane,assume the following:

-   -   GW=5000 lbs    -   PROP_Efficiency=80%    -   L/D=20    -   Vtrue=300 fps    -   ROCreqd=1,500 fpm

Solving for the RotorDiameter or wingspan when the engine power for VTOLequals the engine power for climb will result in a preferably balanceddesign in which the wings are utilized as the rotor for rotor flight.

${RotorDiameter} = \frac{\left( {\frac{1}{2}*G\; W} \right)^{1.5}}{\begin{matrix}{\left\lbrack {{PROP\_ Efficiency}*21*\sqrt{\frac{\pi}{4}}} \right\rbrack*} \\{\left\lbrack \frac{{VCruise}*G\; W}{{PROP\_ Efficiency}*550} \right\rbrack\left\lbrack {\frac{\left\lbrack \frac{R\; O\; C_{reqd}}{60} \right\rbrack}{VCruise} + \left( \frac{L}{D} \right)^{- 1}} \right\rbrack}\end{matrix}}$

In this example only, RotorDiameter=wingspan=68.7 ft.

The previous calculations matched engine power provided by a propellerfor vertical and hovering flight and fixed wing flight climb. However,to eliminate the need for mechanical gearing between the flight modes,the engine is desirably secured laterally on the wing to provide thedesired rotor torque at the rotor RPM.

Assuming the aircraft when it is in fixed wing configuration has anaspect ratio (AR) of 20 the RPM and torque required may be determined.

Near an advance ratio of zero (hover) an AR=20 wing has theseproperties.

RotorThrust Coefficient, CT=0.194

${Thrust} = {\left( {C\; T*\rho*\frac{R\; P\; M}{60}} \right)^{2}*{RotorDiameter}^{4}}$${1.2*G\; W} = {\left( {C\; T*\rho*\frac{R\; P\; M}{60}} \right)^{2}*{RotorDiameter}^{4}}$

Solving for the rotor rotations per minute results in 46 rpm for thewings when they act as a rotor. Recall:

${VTOL\_ SHP}_{reqd} \cong \frac{\left( {1.2*G\; W} \right)^{1.5}}{\left\lbrack {21*{RotorDiameter}*\sqrt{\frac{\pi}{4}}} \right\rbrack}$

Thus VTOL_SHP_(reqd)=454.7 hp.

Therefore:

${Torque} = \frac{{VTOL\_ SHP}*550}{2*\pi*\frac{R\; P\; M}{60}}$

and Torque=41,623 ft-lbs.

Assuming the thrust of the engines in VTOL or vertical/hovering flightis defined as:

${Thrust} = \frac{{PROP\_ Efficiency}*S\; H\; P*550}{V@{prop}}$

Where V@prop is the relative wind at the engine station on the rotatingwing, given by:

${{V@{prop}} = {\left( \frac{R\; P\; M}{60} \right)*\pi*{Engine}\mspace{14mu}\%\mspace{14mu}{Semispan}*{RotorDiameter}}}\;$

Then V@prop=82.7 fps.

For engines secured at 50% semispan the available thrust is:

${{Total\_ Thrust}{\_ Avail}} = \frac{{PROP}_{Efficiency}*{SHP}*550}{\left( {V@{prop}} \right)}$

Solving the equation results in Total Thrust Available=2,425 lbs.

From the Rotor Torque Equation:Torque=TotalThrust_(reqd) *Y

Rearranged:

${TotalThrust}_{reqd} = \frac{Torque}{Y}$

Since the rotor diameter, or total wingspan, is 68.7 ft, as calculatedabove for this example only, an engine located at 50% semi-span has alever arm (Y) of 17.16 ft.

Therefore, in this example, the Total Thrust Required is 2,425 lbs,which equals the Total Thrust Available as calculated above.

The equivalence of the Total Thrust Available and the Total ThrustRequired illustrates that for this example, a balanced design wasachieved without needing a gearbox.

As a further example, a balanced 5,000 lb aircraft such as thatdescribed in FIGS. 1-10 desirably has a wingspan of 68.7 ft, with a pairof 228 hp engines rigged 7.5 ft from the centerline. A similar analysisfor a 500 lb aircraft with an AR=25, Vcruise=100 fps,Prop_Efficiency=65%, L/D=15, Ythrust/Semispan=40% and ROC=1000 fpmyields a rotor diameter of 37.2 ft, powered by 2 16.3 Hp engines rigged7.45 ft from centerline. In other embodiments, a balanced aircraft mayweigh between 400 lb and 100,000 lb, between 1,000 lb and 50,000 lb, orbetween 2,500 lb and 10,000 lb. In other embodiments, the wingspan ofthe aircraft may be between 6 ft and 250 ft, between 20 ft and 150 ft,or between 50 ft and 100 ft. In some embodiments, the weight of theaircraft is less than 20,000 lb, less than 10,000 lb, less than 5,000lb, less than 3,000 lb, or less than 1,000 lb. In other embodiments, theengines may be rigged between 2 ft and 50 ft from the centerline,between 4 ft and 25 ft from the centerline, or between 6 ft and 15 ftfrom the centerline. In some embodiments, the wingspan of the aircraftis less than 50 ft, less than 20 ft, or less than 10 ft. However, thepresent invention is not limited to aircraft of any specific weight orwingspan or engines of a particular power output.

Although this application discloses certain preferred embodiments andexamples, it will be understood by those skilled in the art that thepresent inventions extend beyond the specifically disclosed embodimentsto other alternative embodiments and/or uses of the invention andobvious modifications and equivalents thereof. For example, while thepreferred embodiment is an unmanned aircraft, the aircraft could bemanned with a cockpit desirably located in nose section 102. Further,the various features of these inventions can be used alone, or incombination with other features of these inventions other than asexpressly described above. While the disclosed embodiments are primarilydirected to an aircraft capable of fixed wing and rotor flight modes,aspects of the invention may also be used in connection with other typesof aircraft. Thus, it is intended that the scope of the presentinventions herein disclosed should not be limited by the particulardisclosed embodiments described above, but should be determined only bya fair reading of the claims that follow

What is claimed is:
 1. An aircraft capable of fixed wing and rotorflight modes, comprising: a fuselage main body defining a longitudinalaxis, the fuselage main body having a nose, a tail, and a rotatable wingattachment region disposed between the nose section and the tailsection; a plurality of dual-purpose wings, including a first wing and asecond wing, rotatably mounted to said wing attachment region of saidfuselage main body for a fixed wing flight mode and for a rotor flightmode, in which the fixed wing flight mode is defined as flight in whichsaid wings are maintained rotationally stationary relative to thelongitudinal axis and the rotor flight mode is defined as flight inwhich said wings rotate about the longitudinal axis; and a plurality ofengines secured to said wings, including a first engine secured to saidfirst wing in an intermediate region of said first wing and a secondengine secured to said second wing in an intermediate region of saidsecond wing, in which the plurality of engines are each solely confinedto the wings, said first engine and said second engine includingpropellers.
 2. The aircraft of claim 1, wherein each wing includes arear swept portion proximate to the fuselage and a forward swept portiondistal to the rear swept portion.
 3. The aircraft of claim 1, whereineach of the plurality of engines are secured to a corresponding wing ofthe plurality of wings at a prescribed semi-span distance that isdisposed at or between a first calculated location and a secondcalculated location, in which (a) the first calculated location is alocation along the semi-span calculated to optimize rotor flight modefor the corresponding wing and propulsive power of the correspondingengine, (b) the second calculated location is a location along thesemi-span calculated to optimize fixed wing flight mode for thecorresponding wing and propulsive power of the corresponding engine, and(c) the first calculated location and the second calculated location areno more than 25% apart along of the semi-span distance.
 4. The aircraftof claim 1, further comprising fuel tanks disposed in the wings andoperatively coupled to the plurality of engines.
 5. The aircraft ofclaim 1, wherein the plurality of engines are each secured to said wingsat an equalizing position along the semi-span of each wing.
 6. Theaircraft of claim 1, further comprising a plurality of tail stabilizingsurfaces.
 7. The aircraft of claim 6, wherein said plurality of tailstabilizing surfaces further comprise at least one control surface forcontrolling the orientation of the aircraft when the aircraft is flyingin fixed wing mode or in rotor flight mode.
 8. The aircraft of claim 1,wherein the plurality of dual-purpose wings consist of the first wingand the second wing; and the plurality of engines consist of the firstengine and the second engine.
 9. An aircraft capable of fixed wing androtor flight modes, comprising: a fuselage main body defining alongitudinal axis, the fuselage main body having a nose, a tail, and awing attachment region disposed between the nose section and the tailsection; a plurality of dual-purpose wings, including a first wing and asecond wing, rotatably mounted to said wing attachment region of saidfuselage main body for a fixed wing flight mode and for a rotor flightmode, in which the fixed wing flight mode is defined as flight in whichsaid wings are maintained rotationally stationary relative to thelongitudinal axis and the rotor flight mode is defined as flight inwhich said wings rotate about the longitudinal axis; a plurality ofengines secured to said wings, including a first engine secured to saidfirst wing in an intermediate region of said first wing and a secondengine secured to said second wing in an intermediate region of saidsecond wing, in which the plurality of engines are each solely confinedto the wings; and a plurality of fuel tanks disposed in the plurality ofdual-purpose wings and operatively coupled to the plurality of engines.10. The aircraft of claim 9, wherein the plurality of engines are eachsecured to said wings at an equalizing position along the semi-span ofeach wing.
 11. The aircraft of claim 9, wherein the plurality ofdual-purpose wings consist of the first wing and the second wing; andthe plurality of engines consist of the first engine and the secondengine.
 12. The aircraft of claim 9, wherein each wing includes aproximal wing portion and a distal wing portion, having an engine of theplurality of engines secured between said proximal wing portion and saiddistal wing portion.
 13. The aircraft of claim 9, further comprising aplurality of tail stabilizing surfaces.
 14. The aircraft of claim 13,wherein said plurality of tail stabilizing surfaces further comprise atleast one control surface for controlling the orientation of theaircraft when the aircraft is flying in fixed wing mode or in rotorflight mode.
 15. An aircraft capable of fixed wing and rotor flightmodes, comprising: a fuselage main body defining a longitudinal axis,the fuselage main body having a nose, a tail, and a wing attachmentregion disposed between the nose section and the tail section; aplurality of dual-purpose wings, including a first wing and a secondwing, rotatably mounted to said wing attachment region of said fuselagemain body for a fixed wing flight mode and for a rotor flight mode, inwhich the fixed wing flight mode is defined as flight in which saidwings are maintained rotationally stationary relative to thelongitudinal axis and the rotor flight mode is defined as flight inwhich said wings rotate about the longitudinal axis, wherein each wingincludes a rear swept portion proximate to the fuselage and a forwardswept portion distal to the rear swept portion; a plurality of enginessecured to said wings, including a first engine secured to said firstwing in an intermediate region of said first wing and a second enginesecured to said second wing in an intermediate region of said secondwing, in which the plurality of engines are each solely confined to thewings.
 16. The aircraft of claim 15, wherein the first engine isdisposed between the rear swept portion and the forward swept portion ofthe first wing, and the second engine is disposed between the rear sweptportion and the forward swept portion of the second wing.
 17. Theaircraft of claim 15, wherein the plurality of engines are each securedto said wings at an equalizing position along the semi-span of eachwing.
 18. The aircraft of claim 15, wherein the plurality ofdual-purpose wings consist of the first wing and the second wing; andthe plurality of engines consist of the first engine and the secondengine.
 19. The aircraft of claim 15, further comprising a plurality oftail stabilizing surfaces.
 20. The aircraft of claim 19, wherein saidplurality of tail stabilizing surfaces further comprise at least onecontrol surface for controlling the orientation of the aircraft when theaircraft is flying in fixed wing mode or in rotor flight mode.